1. Field of the Invention
The present invention relates generally to aluminum alloys, and in particular, to such alloys useful in the aerospace industry suitable for use in fuselage applications.
2. Description of Related Art
In today's civil aircraft industry, and in particular for fuselage applications, there is a strong incentive to reduce both weight and cost. The fuselage of a commercial transport aircraft is subject to a complex set of loads, depending on the phase of operation (take-off, cruise, maneuvering, landing . . . ) and environmental conditions (gusts, headwinds, . . . ). Furthermore, different parts of the fuselage are subject to different loadings. In spite of this complexity, it is possible to distinguish major design selection criteria that determine the weight of the structure, some impacting total weight more than others.
For example, compression and shear-compression resistance are extremely important design criteria, since the heaviest fuselage shells are loaded by compression. In order for a new material to allow weight reductions of these compressively loaded shells, this new material should have high Young's modulus, high 0.2% proof stress (to resist buckling) and low density.
A second important design criterion is residual strength of longitudinally cracked shells. Aircraft certification regulations require damage tolerant design, so it is common practice to consider large longitudinal or circumferential cracks in fuselage shells, proving that a certain level of tension can be applied without catastrophic fracture. One known material property governing design here is the plane stress fracture toughness. Any single critical stress intensity factor, however, provides only a limited view of fracture toughness. The development of an R-Curve is a widely recognized method to characterize fracture toughness properties. The R-curve represents the evolution of the stress intensity factor for crack growth as a function of crack extension, under monotonic loading. The R-curve enables the determination of the critical load for unstable fracture for any configuration relevant to cracked aircraft structures. The values of stress intensity factor and crack extension are effective values as defined by ASTM E561. The length of the R-curve—i.e. maximum crack extension of the curve—is an important parameter in itself for fuselage design. The generally employed analysis of conventional tests on center cracked panels gives an apparent stress intensity factor at fracture [KC0]. KC0 does not vary significantly as a function of R-curve length, especially when the R-curve slope is close to the slope of the curve relating the applied stress intensity factor to the crack length (applied curve). However in a real airframe structure such as a panel with attached stiffeners, when a crack progresses under a non-broken stiffener, the applied curve drops due to the bridging effect of the stiffener. In this case a local minimum of the applied curve can occur for a crack length larger than the initial considered crack length plus crack extension under monotonic loading. As such, larger loads at unstable fracture are then allowed for long R-curves. It is thus of interest to have longer R-curve, even for identical conventionally determined critical stress intensity factors.
For products with identical mechanical properties, lower density is clearly beneficial for air frame weight. A third important design criterion is thus material density. Moreover, large parts of the fuselage are not so heavily loaded and the weight of the design is limited by a certain limit generally called “minimum gauge”. The concept of minimum gauge corresponds to the thinnest gauge practicable for manufacturing (particularly handling of panels) and repair (patch riveting). The only way to reduce weight in minimum gauge design is to use a lower density material.
Other important factors affecting material selection include propagation of cracks under fatigue loading, either under constant amplitude loading or with variable amplitude (because of maneuvers and gusts, especially in the longitudinal direction, but also around the wing, in all directions).
Currently, the fuselages of civil aircraft are for the most part made from 2024, 2056, 2524, 6013, 6156 or 7475 alloy sheet or thin plates, clad on either surface with a low composition aluminum alloy, such as a 1050 or 1070 alloy, for example. The purpose of the cladding alloy is to provide sufficient corrosion resistance. Slightly generalized or pitting corrosion is tolerable, but corrosion must not penetrate to attack the core alloy. There is a trend to try using unclad materials for fuselage design, for cost reduction. Corrosion resistance, and particularly resistance to intergranular corrosion and stress corrosion cracking is thus an important aspect of properties of suitable fuselage panels.
As stated above, the only way to reduce weight in some cases is to reduce the density of the materials used for construction of the aircraft. Aluminum-lithium alloys have long been recognized as an effective solution to reduce weight because of the low density of these alloys. However, the different requirements cited above, namely, having a high Young modulus, high compression resistance, high damage tolerance and high corrosion resistance, have not been met simultaneously by prior art aluminum-lithium alloys. In particular, obtaining a high fracture toughness with these alloys has proven to be difficult. Prasad et al, for example, state recently (Sadhana, vol. 28, Parts 1&2, February/April 2003 pp. 209-246) that “Al—Li alloys are prime candidate materials to replace traditionally used Al alloys. Despite their numerous property advantages, low tensile ductility and inadequate fracture toughness, especially in the through thickness-directions, militates against their acceptability”. Today, Al—Li alloys have been limited to very specific military applications such as high temperature resistance materials, improved cryogenic fracture toughness materials for aerospace applications, and certain parts in helicopters and military aircraft fuselage parts.
U.S. Pat. No. 5,032,359 (Martin Marietta) describes a family of alloys based upon aluminum-copper-magnesium-silver alloys to which lithium has been added, within specific ranges and which exhibit superior ambient- and elevated-temperature strength, superior ductility at ambient and elevated temperatures, extrudability, forgeability, weldability, and an unexpected natural aging response. The examples describe extruded products. No information is provided on toughness, resistance to fatigue crack or resistance to corrosion. In a preferred embodiment, the alloy includes an aluminum base metal, from 3.0 to 6.5% of copper, from 0.05 to 2.0% of magnesium, from 0.05 to 1.2% of silver, from 0.2 to 3.1% of lithium, from 0.05 to 0.5% of a grain refiner selected from zirconium, chromium, manganese, titanium, boron, hafnium, vanadium, titanium diboride, and mixtures thereof.
U.S. Pat. No. 5,122,339 (Martin Marietta) is a continuation in part of the '359 patent mentioned supra. It additionally discloses the use of similar alloys as welding alloys or weld alloys.
U.S. Pat. No. 5,211,910 (Martin Marietta) describes aluminum-base alloys containing Cu, Li, Zn, Mg and Ag which possess highly desirable properties, such as relatively low density, high modulus, high strength/ductility combinations, strong natural aging response with and without prior cold work, and high artificially aged strength with and without prior cold work. The alloys may comprise from about 1 to about 7 weight percent Cu, from about 0.1 to about 4 weight percent Li, from about 0.01 to about 4 weight percent Zn, from about 0.05 to about 3 weight percent Mg, from about 0.01 to about 2 weight percent Ag, from about 0.01 to about 2 weight percent grain refiner selected from Zr, Cr, Mn, Ti, Hf, V, Nb, B and TiB2, and the balance Al along with incidental impurities. The '910 patent discloses how Zn additions may be used to reduce the levels of Ag present in the alloys taught in U.S. Pat. No. 5,032,359, in order to reduce cost.
U.S. Pat. No. 5,455,003 (Martin Marietta) discloses a method for the production of aluminum-copper-lithium alloys that exhibit improved strength and fracture toughness at cryogenic temperatures. Improved cryogenic properties are achieved by controlling the composition of the alloy, along with processing parameters such as the amount of cold-work and artificial aging. The product is used for cryogenic tanks in space launch vehicles.
U.S. Pat. No. 5,389,165 (Reynolds) discloses an aluminum-based alloy useful in aircraft and aerospace structures which has low density, high strength and high fracture toughness of the following formula: CuaLibMgcAgdZreAlbal wherein a, b, c, d, e and bal indicate the amount in wt. % of alloying components, and wherein 2.8<a<3.8, 0.80<b<1.3, 0.20<c<1.00, 0.20<d<1.00 and 0.08<e<0.46. Preferably, the copper and lithium components are controlled such that the combined copper and lithium content is kept below the solubility limit to avoid loss of fracture toughness during elevated temperature exposure. The relationship between the copper and lithium contents also should meet the following relationship: Cu(wt. %)+1.5 Li(wt. %)<5.4. Special stretching conditions, between 5 and 11% have been applied. Examples are limited to a thickness of 19 mm and zirconium content superior or equal to 0.13 wt %.
US 2004/0071586 (Alcoa) discloses an Al—Cu—Mg alloy including from 3 to 5 weight percent Cu, from 0.5 to 2 weight percent Mg and from 0.01 to 0.9 weight percent Li. According to this application, toughness properties of alloys having additions of from 0.2 to 0.7 weight percent Li are significantly improved compared to similar alloys containing either no Li or a greater amount of Li.
There is a need for a high strength, high fracture toughness, and especially high crack extension before unstable fracture, high corrosion resistance Al—Li alloy for aircraft applications, and in particular for fuselage sheet applications.